For many years it has been generally acknowledged throughout the gas turbine industry that the lowest cost engine configuration is that utilizing a single-stage radial compressor and turbine mounted back-to-back on a common overhung, cantilevered shaft.
Continued development over the years of this engine arrangement has led to increased pressure ratios and turbine inlet temperatures to as high as 6.0:1 and 1,700.degree. F, respectively.
In order to increase operating turbine inlet temperatures beyond 1,700.degree. F, methods of cooling the turbine rotor were required. After studying many cooling methods, it was determined that a monorotor arrangement in which the two rotors were integrated and the hot section conduction cooled by the cold compressor heat sink offered substantial advantages.
Several low-temperature, low pressure ratio monorotor gas turbines were developed during the 1950's. However, the development of small single-stage centrifugal compressors and radial inflow turbines has been guided primarily by the priority of low manufacturing cost rather than operational efficiency or maximum performance. In recent years, however, the need for optimum operational efficiencies and/or maximum performance has continued to become an increasingly greater factor in the gas turbine engine field. Projections concerning the availability of gas, oil and other hydro carbon fluids indicate an ever-increasing demand for an ever-decreasing supply. This, of course, means that the relative value of this factor is, and will continue to be, an important element in the gas turbine engine selection process.
Increased cycle efficiencies can be obtained by increasing turbine inlet temperature but significant consideration to the maximum allowable metal temperature of the turbine rotor is necessary. In general, the life expectancy of a high temperature turbine stage is determined by the peak metal temperature and distribution, which, of course, govern blading oxidation and thermal fatigue characteristics.
Stress rupture life is also dependent upon the temperature of the metal. Of primary importance is the allowable metal temperature of the monorotor. High-strength, high-temperature metals are now available that are capable of withstanding rotor metal temperatures of about 1,500.degree. F. Higher metal temperatures may be tolerated in the lower stressed turbine nozzle where metal temperatures of 1,800.degree. F are acceptacle. Consequently, if the monorotor metal temperature is maintained below 1,500.degree. F, no additional cooling will be required.
Further, it is obvious that the temperature within the monorotor will not be entirely uniform. Axial and radial temperature gradients will exist due to:
1. The finite width of the disc separating the compressor and turbine; PA1 2. The height and taper ratio of the turbine and compressor blades; PA1 3. The existance and use of additional heat sinks, such as the shaft; and PA1 4. The effects of combustor exit temperature variations.
The present invention as herein and hereinafterwards disclosed, in the main, overcomes many of the problems mentioned hereinbefore of the monorotor gas turbine engine designs by providing a uniquely structured gas turbine utilizing a novel turbine rotor cooling technique.